Composites are used because of their high strength to weight ratio. Fastening composite parts, however, is often troublesome. For example, in attaching an aircraft structural skin member to a supporting stringer, bolts and/or rivets may be used, but such fasteners add weight, increase fabrication costs, and can contribute to local failure modes between the plies of the laminate composite. Brazing is also used to join two composite parts, (see U.S. Pat. No. 5,021,107), but does not always provide sufficient strength to maintain the integrity of the joint between the two composites under load.
It was discovered by the applicant that composite structures can be reinforced to prevent delamination and cracking by inserting pins such as boron fibers through the thickness of the composite. The pins are inserted into a thermally decompressable foam body which is then placed on the composite. The application of increased temperature and pressure in an autoclave drives the pins through the thickness of the composite. See, U.S. Pat. No. 4,808,461.
Recently, it was realized by the applicant that the pins can be driven through two composite structures to join them, and that such a structure joined by pins extending through both laminates was stronger than one made by brazing alone and offered significant benefits over rivets or bolts used as fasteners. See, U.S. patent application Ser. No. 08/056,029, and U.S. Pat. No. 5,186,776.
Still, the state of the art lacks a convenient methodology for inserting pins to join composite structures like the skin/stringer arrangement of an aircraft wing since such structures cannot be conveniently placed in an autoclave.